1. Technical Field
This invention relates to gas turbine engines in general, and cooled turbine vane assemblies in particular.
2. Background Information
Turbine sections within an axial flow turbine engine generally include a plurality of rotor assemblies and stationary vane assemblies. The rotor assemblies, each comprising a disk and a plurality of rotor blades circumferentially disposed around the disk, rotate about an axis extending through the turbine engine. The vane assemblies, disposed forward and/or aft of each rotor assembly, include a plurality of vanes and an inner vane support. The outer radial end of each vane is fixed to the casing of the turbine section and the inner radial end of each vane is fixed to the inner vane support.
The thermal environment within the turbine section necessitates cooling both the vane assemblies and the rotor assemblies, although the amount of cooling required depends upon the power setting of the engine. Typically, air at a pressure higher and a temperature lower than that of the core gas flow will be passed through the vane and rotor assemblies as a cooling medium. Bypass air originating from a fan section, or air bled off of a compressor section, provides the cooling air at the higher pressure and lower temperature. In some turbine stages, cooling air passes through the vanes and into the vane support before being directed into an aft rotor assembly by a plurality of tangential on-board injectors (also known as TOBI's). The TOBI's, which are formed within or attached to the inner vane support, direct the air exiting the inner vane support in a direction substantially parallel to the rotor assembly plane of rotation. The injectors are aligned within disk inlet orifices, thereby enabling cooling air exiting the vane support to enter the disk of the rotor assembly and pass thereafter up and into the blades.
A person of skill in the art will recognize that there is a tension between the need to cool the turbine sections and the efficiency of the engine. It is often desirable to cool the turbine sections with air bled from a compressor stage, for example, because it has been worked to a higher pressure and therefore has the energy necessary to travel through a vane assembly and a subsequent rotor assembly. A significant percentage of the work imparted to the air bled from the compressor, however, is lost during the cooling process. The lost work does not add to the thrust of the engine and therefore negatively effects the overall efficiency of the engine.
To minimize the amount of energy lost in the cooling process, it is known to add flow restricting devices either upstream of where cooling air enters each vane, or immediately inside each vane. The flow restricting devices conserve energy by limiting the flow of cooling air into each vane. A consequence of limiting the flow, however, is a pressure drop across the restricting device. The amount of pressure drop permissible is limited because a minimum pressure must be maintained within each vane to prevent the inflow of hot core gas flow, or in the case of a TOBI arrangement the minimum pressure may be that which is necessary to cool the aft rotor assembly. Another consideration in cooling air pressure is the cross-sectional area of apertures within the vanes through which cooling air may exit the vane and pass out into the core gas flow. Theoretically, apertures can be designed which satisfy the cooling needs of the vane and help maintain the minimum pressure necessary within the vane along the leading edge. In reality, however, the cross-sectional area of those optimum apertures is so small that the apertures are either impossible to cast within the vanes, or are cost prohibitive to machine in the vanes. Apertures having a cross-sectional area greater than optimum are therefore utilized, and additional flow is required within the blade to maintain the minimum required pressure. Hence, there is a limit to the extent which cooling air flow can be restricted under presently available vane design.
A further complication in the tension between efficiency and cooling is the future direction of compressor design. For a variety of reasons, it is advantageous to minimize the numbers of compressor stages within a gas turbine engine. Decreasing the number of compressors does not, however, decrease the work requirement of the compressor stages. On the contrary, with fewer number of compressor stages, each stage is required to do more work. Increasing the work of each compressor stage, increases the jump in output gas pressure and temperature between each stage. At present, cooling air is bled off of the compressor stage having an output pressure closest to the minimum required pressure for cooling purposes. Minimizing the difference between the bled gas pressure and the minimum required cooling air pressure, minimizes the amount of work lost to cooling. In future compressor designs, the "jumps" between compressor stages will be greater, thereby potentially increasing the difference in pressure available and that needed for cooling purposes and therefore decreasing the efficiency of the engine.
What is needed, therefore is a turbine vane assembly that may be adequately cooled using a minimum of cooling air, thereby increasing the efficiency of the engine.